This application claims priority to PCT Patent Application No. PCT/US14/071609 filed Dec. 19, 2014 which claims priority to U.S. Patent Application No. 61/918,452 filed Dec. 19, 2013, which are hereby incorporated herein by reference in their entireties.
Gas turbine engines, such as those that power modern commercial and military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust. All sections operating in unison under a wide variety of flight conditions including, but not limited to, idle, take-off, cruise, and deceleration.
The combustor section generally includes a fuel system having a plurality of circumferentially distributed fuel injectors that axially project into a combustion chamber to supply fuel to be mixed with the pressurized air. This fuel supply through the injectors must be controlled by the fuel system to meet the demands of the various flight conditions while maintaining engine efficiency and minimizing emissions. To do so, fuel systems may include complicated and heavy hydromechanical devices that may include a plurality of multi-staged fuel manifolds, isolation or servo valves between individual injectors and the manifolds, low and high pressure fuel pumps delivering and bypassing fuel to the manifolds, and a wide array of electronic control features.
Unfortunately, even with such complex and expensive systems, operation of individual fuel injectors is limited to fuel pressures and flows from the upstream manifolds, and fuel isolation provided by the interposed servo valves. Controlling (ramping up or down) of fuel flow through each individual injector, or individually tuning a fuel injector, to further refine operating performance is generally not available. Yet further, fuel atomization under start-up and low power conditions is limited by available fuel pressure.